1. Field of Invention
This invention relates to voltage regulation, and more particularly to voltage regulation in an electronic engine control system having digital effector type servo actuators.
2. Description of the Prior Art
Modern jet engines, such as gas turbine engines, employ electronic engine control (EEC) units to control the engine operating performance. The EEC units receive signal information on sensed engine operating parameters, such as rotor speed, air intake temperature, and exhaust temperature and pressure levels, and process the information through known signal processing algorithms to provide position signals to each of a plurality of digital effector servo actuators of the type shown and described in a commonly owned, copending application entitled ADAPTIVE CONTROL SYSTEM USING POSITION FEEDBACK, U.S. Ser. No. 586,010 filed on June 1, 1975 by A. N. Martin now U.S. Pat. No. 4,007,361, Feb. 8, 1977, which direct engine performance through position control of a corresponding plurality of controlled surfaces within the engine, such as fuel valves, stator vane positions, and engine exhaust nozzle positions. The EEC units receive AC electrical power from dedicated, engine driven alternators which are directly driven from the engine gearbox, without the speed control provided for the engine driven alternators providing 115 volts rms, 400 hertz general aircraft power. This is due to the lower reliability of such speed controls which prohibit their use in a primary control system, such as that of the EEC unit, where failure of the constant speed device may result in catastrophic engine failure. As a result, the EEC alternators rotate at an angle of velocity equal to the engine RPM levels. Since an alternator having a fixed number of poles provides output voltage signals whose amplitude and frequency are directly proportional to the angular velocity of the rotor, the engine driven EEC alternators provide an output voltage whose amplitude and frequency increases with increasing engine RPM levels.
The general speed range of an aircraft engine varies from ground speeds of 10 to 15 percent of max RPM, to typical cruise speeds in the range of 90 to 100 percent of max RPM, causing an appreciable increase in alternator output electrical power over this range. The EEC units require essentially constant input power (which is independent of engine speed) for proper operation, and since the EEC must provide engine control over the entire engine RPM range, the alternator must be sized to provide the required EEC power at the minimum engine RPM levels. An alternator which is sized to provide the required EEC unit power at 10 percent of max RPM, may provide as much as five times the required EEC power at the cruise speed range of 90 to 100 percent of max RPM, such that at cruise RPM levels there is a significant amount of excess power which is generated by the alternator and which must be dissipated within the EEC unit. In addition to the excess power which must be dissipated, the alternator output voltage amplitude levels must be limited to a determined level to protect the EEC unit components against voltage breakdown, or overstress.
In typical prior art EEC systems, the alternator AC output voltage is rectified to provide an unregulated DC voltage signal, and the unregulated DC signal is presented to a dedicated voltage regulation circuit. The regulation circuit, typically of the series pass type, provides amplitude control of the output signal through feedback control of the current through a series pass transistor. Since the output amplitude is controlled, and the unregulated DC signal increases with engine RPM in dependence on the AC input voltage, the voltage developed across the regulator increases substantially with engine RPM increase resulting in excessive power dissipation within the EEC unit. To eliminate the resulting heat buildup within the unit, complex cooling systems and large surface area heat sinks are required, which in some systems account for as much as 40 percent of the control unit size and weight. This is of particular concern in aircraft engine control systems where both weight and size must be minimized, and in small, portable engine units where the same considerations govern. Furthermore, in view of the necessity of providing improved engines capable of ever higher horsepower to weight ratios to improve fuel economy, the size and weight of the engine control systems must be similarly reduced is such systems are to be practical.